This invention relates to gas turbine engines, and more particularly to cooled turbine elements (e.g., blades and vanes).
Efficiency is limited by turbine element thermal performance. Air from the engine's compressor bypasses the combustor and cools the elements, allowing them to be exposed to temperatures well in excess of the melting point of the element's alloy substrate. The cooling bypass represents a loss and it is therefore desirable to use as little air as possible. Trailing edge cooling of the element's airfoil is particularly significant. Aerodynamically, it is desirable that the trailing edge portion be thin and have a low wedge angle to minimize shock losses.
In one common method of manufacture, the main passageways of a cooling network within the element airfoil are formed utilizing a sacrificial core during the element casting process. The airfoil surface may be provided with holes communicating with the network. Some or all of these holes may be drilled. These may include film holes on pressure and suction side surfaces and holes along or near the trailing edge.